1. Field of the Invention
This invention relates to cooled turbine blades and more particularly to a transpiration cooled blade having a ceramic airfoil portion.
2. Description of the Prior Art
Cooled turbine blades are well known in the art. One means of blade cooling offering great potential is referred to as transpiration cooling and is accomplished by introducing a cooling fluid into a hollow airfoil portion of the blade, with the skin of the airfoil portion being porous through minute passages for the effusion of the fluid therethrough. This cools the blade by transporting the heat within the blade to the fluid and further, the fluid provides a boundary layer on the exterior of the blade surface preventing the hot motive gases from direct contact therewith. As effective as such cooling is however, in the projected range of temperatures of operation necessary to obtain 50 to 55% efficiency for a gas turbine engine (turbine inlet temperatures must then approach 2500.degree. to 3000.degree. F.) the high temperature alloys from which most blades are fabricated tend to oxidize and the minute transpiration flow paths thus become plugged.
In view of the above, the use of ceramic blades is actively being investigated. However, ceramic (i.e. Si.sub.3 N.sub.4 and SiC) have limited strength in tension and also tend to glassify at faults and erode at high temperature. Therefore, even though the ceramics permit a higher turbine inlet temperature, it would be preferable to provide cooling to such ceramic blades to reduce the probability of their failure at these temperatures. Thus, transpiration cooled ceramic blades offer a solution to permitting a turbine inlet temperature in the range of 3000.degree. F. One such blade is disclosed in U.S. Pat. No. 3,240,468 wherein, to relieve internal stress due to thermal gradients across the blade, a different amount of cooling fluid is directed to separate portions of the cooled blade. Further, in that the present invention involves a blade assembled from a plurality of stacked washers forming the airfoil portion of the blade, U.S. Pat. Nos. 3,301,526 and 3,515,499 are relevant for showing a prior art turbine vane comprising a plurality of airfoil-shaped wafers stacked to form a cooled vane.